Rotor assembly

ABSTRACT

The present disclosure relates to a rotor assembly for a gas turbine engine, the rotor assembly comprising a first rotor stage having a first disc portion with a peripheral first rim portion and a second rotor stage, the second rotor stage having a second disc portion with a peripheral second rim portion. The second rotor stage is axially adjacent and downstream of the first rotor stage and the second rim portion has an axial extension extending towards the first rim portion such that the axial extension of the second rim portion defines a rotor drum cavity between the first and second disc portions. The second rotor stage further comprises a drive arm extending within the drum cavity to the first disc portion, the drive arm being connected to the first disc portion by at least one connector. The drive arm divides the drum cavity into radially outer rim cavity portion and a radially inner main cavity portion. The rotor assembly further comprises a rim seal located between the axial extension of the second rim portion and the first rim portion, and a pressure equalisation path extending from the rim cavity portion to the main cavity portion.

CROSS-REFERENCE TO RELATED APPLICATIONS

This specification is based upon and claims the benefit of priority fromUK Patent Application Number GB2001689.5 filed on 7^(th) of Feb. 2020,the entire contents of which are incorporated herein by reference.

BACKGROUND Technical Field

The present disclosure relates to a rotor assembly, in particular acompressor assembly such as a high-pressure compressor assembly havingmultiple adjoined and connected rotor/compressor stages.

Description of the Related Art

FIG. 1 illustrates a gas turbine engine 10 having a principal rotationalaxis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23that generates two airflows: a core airflow A and a bypass airflow B.The gas turbine engine 10 comprises a core 11 that receives the coreairflow A. The engine core 11 comprises, in axial flow series, alow-pressure compressor 14, a high-pressure compressor 15, combustionequipment 16, a high-pressure turbine 17, a low-pressure turbine 19 anda core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. Thebypass airflow B flows through the bypass duct 22. The fan 23 isattached to and driven by the low-pressure turbine 19 via a shaft 26 andan epicyclic gearbox 30.

In use, the core airflow A is accelerated and compressed by thelow-pressure compressor 14 and directed into the high-pressurecompressor 15 where further compression takes place. The compressed airexhausted from the high-pressure compressor 15 is directed into thecombustion equipment 16 where it is mixed with fuel and the mixture iscombusted. The resultant hot combustion products then expand through,and thereby drive, the high pressure and low-pressure turbines 17, 19before being exhausted through the nozzle 20 to provide some propulsivethrust. The high-pressure turbine 17 drives the high-pressure compressor15 by a suitable interconnecting shaft 27. The fan 23 generally providesthe majority of the propulsive thrust. The epicyclic gearbox 30 is areduction gearbox.

As shown in FIG. 4 , each compressor 14, 15 comprises multiple stagesextending between an upstream end (proximal the fan) and a downstreamend (distal the fan).

Each stage includes a rotor disc 100 a, 100 b having a peripheral rim101 a, 101 b and multiple circumferentially arranged blades 102 a, 102 bthat are either integral with the rim 101 a, 101 b of the rotor disc 100a, 100 b or affixed to the rim 101 a, 101 b of the rotor disc 100 a, 100b, followed by multiple stator vanes 103 attached to a stationary ring(not shown). The stator vanes may be shrouded or cantilevered.

Each rotor disc 100 a, 100 b includes a respective axial extension 104a, 104 b, the axial extensions defining a drum cavity 105. In order toconnect adjacent stages, a drive arm 106 is provided which extends fromthe axial extension 104 b towards the rotor disc 104 a of the adjacentstage. The free end of the drive arm 106 is connected to the adjacentrotor disc 104 a by a plurality of circumferentially arranged bolts ortie bars 107.

The drive arm 106 section divides the drum cavity into a radially outer(rim) cavity portion 108 and a radially inner (main) cavity portion 109.The rim cavity portion 108 is at the pressure of the gas flow path (inwhich the blades 102 a, 102 b and vanes 103 are mounted). This meansthat the bolts/tie bars 107 not only have to maintain the structuralintegrity of the connection between the stages (under high axial loads),they also have to seal the gas path. The pressure in the rim cavityportion 108 exerts a significant proportion (20-30%) of the total axialload on the bolts/tie bars 107.

The high axial loads on the bolts/tie bars 107 mean that a high numberof bolts/tie bars 107 have to be used which increases the weight of thecompressor and/or highest strength bolt/tie bar material has to be usedwhich increases the cost of the compressor.

SUMMARY

According to a first aspect there is provided a rotor assembly for a gasturbine engine, the rotor assembly comprising:

a first rotor stage having a first disc portion with a peripheral firstrim portion;

a second rotor stage, the second rotor stage having a second discportion with a peripheral second rim portion,

wherein the second rotor stage is axially adjacent and downstream of thefirst rotor stage and the second rim portion has an axial extensionextending towards the first rim portion such that the axial extension ofthe second rim portion defines a rotor drum cavity between the first andsecond disc portions,

wherein the second rotor stage further comprises a drive arm extendingwithin the drum cavity to the first disc portion, the drive arm beingconnected to the first disc portion by at least one connector, whereinthe drive arm divides the drum cavity into radially outer rim cavityportion and a radially inner main cavity portion,

wherein the rotor assembly further comprises a rim seal located betweenthe axial extension of the second rim portion and the first rim portion,

wherein the rotor assembly comprises a pressure equalisation pathextending from the rim cavity portion to the main cavity portion, and

wherein the drive arm comprises an oblique portion extending from aradially inner surface of the axial extension of the second rim portion.

By providing a rim seal between the rim portions and a pressureequalisation path between the rim cavity portion and the main cavityportion, the pressure within the rim cavity portion substantiallymatches the pressure in the main cavity portion rather than the pressureof the gas path (which is radially outwards of the rim portions).Accordingly, the load applied to the connector(s) is significantlyreduced because the lower pressure in the rim cavity no longer imparts ahigh axial separation pressure load. The connector(s) only have toprovide the structural connection between the two rotor stages and nolonger have to provide the seal against the pressure in the gas path.This increases the life of the connector(s). It may also reduce thecosts and weight associated with the rotor assembly since the number ofconnectors can be reduced and/or reduced strength (and therefore reducedcost) connector material may be used.

In some embodiments, the drive arm extends from a radially inner surfaceof the axial extension of the second rim portion. In some embodiments,the drive arm comprises an oblique portion extending obliquely bothaxially and radially inwards e.g. axially and radially inwards from theradially inner surface of the axial extension.

In some embodiments, the drive arm comprises a radial portion extendingfrom the oblique portion, the radial portion lying in abutment with thefirst disc portion and wherein the connector(s) extend through the firstdisc portion and the radial portion of the drive arm.

In some embodiments, the pressure equalisation path is provided throughthe drive arm. The pressure equalisation path may be provided throughthe oblique portion of the drive arm. For example, the oblique portionof the drive arm may comprise at least one aperture extending from aradially outer surface (in the rim cavity portion) to a radially innersurface (in the main cavity portion).

The rim seal at least partly seals the rim cavity portion from the gasflow path (which is radially outwards of the rim portions). In someembodiments, the rim seal is adapted to completely seal the rim cavityportion from the gas path. In other embodiments, the rim seal may beadapted allow a controllable leakage.

In some embodiments, the rim seal may comprise a folded sheet seal (e.g.a folded metal sheet seal) having a first leaf in abutment with thefirst rim portion and a second leaf in abutment with the axial extensionof the second rim portion. In some embodiments, the axial extension mayhave a chamfered edge proximal the first rim portion. In theseembodiments, the second leaf may lie in abutment with the chamferededge.

In some embodiments, the rim seal may comprise an interference seal.

In some embodiments, there is a plurality of connectors, the connectorsbeing circumferentially spaced around the radial portion of the drivearm/first disc portion.

The or at least one of the connectors may be a bolt. The or at least oneconnection element may be a tie bar or any other suitable connector.

The peripheral rim portions are for receiving the rotor aerofoils. Aplurality of aerofoils will be circumferentially arranged around the rimportion. They may be affixed to or integral with the rim portion.

Each rotor stage will comprise a series of circumferentially arrangedstators axially downstream of the aerofoil. The stators are preferablycantilevered stators. The stators associated with the first rotor stagewill be radially aligned and radially outwards of the axial extension ofthe second rim portion.

In some embodiments, the rotor assembly is a compressor assembly e.g. ahigh-pressure compressor assembly and the first rotor stage is a firstcompressor stage and the second rotor stage is a second (adjacent)compressor stage.

In a second aspect, there is provided a gas turbine engine comprising arotor assembly or a compressor assembly (e.g. a high-pressure compressorassembly) according to the first aspect.

Accordingly, the present disclosure may relate to a gas turbine engine.Such a gas turbine engine may comprise an engine core comprising aturbine, a combustor, a compressor, and a core shaft connecting theturbine to the compressor. Such a gas turbine engine may comprise a fan(having fan blades) located upstream of the engine core.

Arrangements of the present disclosure may be particularly, although notexclusively, beneficial for fans that are driven via a gearbox.Accordingly, the gas turbine engine may comprise a gearbox that receivesan input from the core shaft and outputs drive to the fan so as to drivethe fan at a lower rotational speed than the core shaft. The input tothe gearbox may be directly from the core shaft, or indirectly from thecore shaft, for example via a spur shaft and/or gear. The core shaft mayrigidly connect the turbine and the compressor, such that the turbineand compressor rotate at the same speed (with the fan rotating at alower speed).

The gas turbine engine as described and/or claimed herein may have anysuitable general architecture. For example, the gas turbine engine mayhave any desired number of shafts that connect turbines and compressors,for example one, two or three shafts. Purely by way of example, theturbine connected to the core shaft may be a first turbine, thecompressor connected to the core shaft may be a first compressor, andthe core shaft may be a first core shaft. The engine core may furthercomprise a second turbine, a second compressor, and a second core shaftconnecting the second turbine to the second compressor. The secondturbine, second compressor, and second core shaft may be arranged torotate at a higher rotational so speed than the first core shaft.

In such an arrangement, the second compressor may be positioned axiallydownstream of the first compressor. The second compressor may bearranged to receive (for example directly receive, for example via agenerally annular duct) flow from the first compressor.

The gearbox may be arranged to be driven by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example the first core shaft in the example above). For example,the gearbox may be arranged to be driven only by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example only be the first core shaft, and not the second coreshaft, in the example above). Alternatively, the gearbox may be arrangedto be driven by any one or more shafts, for example the first and/orsecond shafts in the example above.

In any gas turbine engine as described and/or claimed herein, acombustor may be provided axially downstream of the fan andcompressor(s). For example, the combustor may be directly downstream of(for example at the exit of) the second compressor, where a secondcompressor is provided. By way of further example, the flow at the exitto the combustor may be provided to the inlet of the second turbine,where a second turbine is provided. The combustor may be providedupstream of the turbine(s).

The or each compressor (for example the first compressor and secondcompressor as described above) may comprise any number of stages, forexample multiple stages. Each stage may comprise a row of rotor bladesand a row of stator vanes, which may be variable stator vanes (in thattheir angle of incidence may be variable). The row of rotor blades andthe row of stator vanes may be axially offset from each other.

The or each turbine (for example the first turbine and second turbine asdescribed above) may comprise any number of stages, for example multiplestages. Each stage may comprise a row of rotor blades and a row ofstator vanes. The row of rotor blades and the row of stator vanes may beaxially offset from each other.

Each fan blade may be defined as having a radial span extending from aroot (or hub) at a radially inner gas-washed location, or 0% spanposition, to a tip at a 100% span position. The ratio of the radius ofthe fan blade at the hub to the radius of the fan blade at the tip maybe less than (or on the order of) any of: 0.4, 0.39, 0.38 0.37, 0.36,0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. Theratio of the radius of the fan blade at the hub to the radius of the fanblade at the tip may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds). These ratios may commonly be referred to as the hub-to-tipratio. The radius at the hub and the radius at the tip may both bemeasured at the leading edge (or axially forwardmost) part of the blade.The hub-to-tip ratio refers, of course, to the gas-washed portion of thefan blade, i.e. the portion radially outside any platform.

The radius of the fan may be measured between the engine centreline andthe tip of a fan blade at its leading edge. The fan diameter (which maysimply be twice the radius of the fan) may be greater than (or on theorder of) any of: 250 cm (around 100 inches), 260 cm, 270 cm (around 105inches), 280 cm (around 110 inches), 290 cm (around 115 inches), 300 cm(around 120 inches), 310 cm, 320 cm (around 125 inches), 330 cm (around130 inches), 340 cm (around 135 inches), 350 cm, 360 cm (around 140inches), 370 cm (around 145 inches), 380 (around 150 inches) cm or 390cm (around 155 inches). The fan diameter may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds).

The rotational speed of the fan may vary in use. Generally, therotational speed is lower for fans with a higher diameter. Purely by wayof non-limitative example, the rotational speed of the fan at cruiseconditions may be less than 2500 rpm, for example less than 2300 rpm.Purely by way of further non-limitative example, the rotational speed ofthe fan at cruise conditions for an engine having a fan diameter in therange of from 250 cm to 300 cm (for example 250 cm to 280 cm) may be inthe range of from 1700 rpm to 2500 rpm, for example in the range of from1800 rpm to 2300 rpm, for example in the range of from 1900 rpm to 2100rpm. Purely by way of further non-limitative example, the rotationalspeed of the fan at cruise conditions for an engine having a fandiameter in the range of from 320 cm to 380 cm may be in the range offrom 1200 rpm to 2000 rpm, for example in the range of from 1300 rpm to1800 rpm, for example in the range of from 1400 rpm to 1600 rpm.

In use of the gas turbine engine, the fan (with associated fan blades)rotates about a rotational axis. This rotation results in the tip of thefan blade moving with a velocity U_(tip). The work done by the fanblades 13 on the flow results in an enthalpy rise dH of the flow. A fantip loading may be defined as dH/U_(tip) ², where dH is the enthalpyrise (for example the 1-D average enthalpy rise) across the fan andU_(tip) is the (translational) velocity of the fan tip, for example atthe leading edge of the tip (which may be defined as fan tip radius atleading edge multiplied by angular speed). The fan tip loading at cruiseconditions may be greater than (or on the order of) any of: 0.3, 0.32,0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (all units in thisparagraph being Jkg⁻¹ K⁻¹/(ms⁻¹)²). The fan tip loading may be in aninclusive range bounded by any two of the values in the previoussentence (i.e. the values may form upper or lower bounds).

Gas turbine engines in accordance with the present disclosure may haveany desired bypass ratio, where the bypass ratio is defined as the ratioof the mass flow rate of the flow through the bypass duct to the massflow rate of the flow through the core at cruise conditions. In somearrangements the bypass ratio may be greater than (or on the order of)any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5,15.5, 16, 16.5, or 17. The bypass ratio may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). The bypass duct may besubstantially annular. The bypass duct may be radially outside the coreengine. The radially outer surface of the bypass duct may be defined bya nacelle and/or a fan case.

The overall pressure ratio of a gas turbine engine as described and/orclaimed herein may be defined as the ratio of the stagnation pressureupstream of the fan to the stagnation pressure at the exit of thehighest pressure compressor (before entry into the combustor). By way ofnon-limitative example, the overall pressure ratio of a gas turbineengine as described and/or claimed herein at cruise may be greater than(or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65,70, 75. The overall pressure ratio may be in an inclusive range boundedby any two of the values in the previous sentence (i.e. the values mayform upper or lower bounds).

Specific thrust of an engine may be defined as the net thrust of theengine divided by the total mass flow through the engine. At cruiseconditions, the specific thrust of an engine described and/or claimedherein may be less than (or on the order of) any of the following: 110Nkg⁻¹s, 105 Nkg⁻¹s, 100 Nkg⁻¹s, 95 Nkg⁻¹s, 90 Nkg⁻¹s, 85 Nkg⁻¹s or 80Nkg⁻¹s. The specific thrust may be in an inclusive range bounded by anytwo of the values in the previous sentence (i.e. the values may formupper or lower bounds). Such engines may be particularly efficient incomparison with conventional gas turbine engines.

A gas turbine engine as described and/or claimed herein may have anydesired maximum thrust. Purely by way of non-limitative example, a gasturbine as described and/or claimed herein may be capable of producing amaximum thrust of at least (or on the order of) any of the following:160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN,450 kN, 500 kN, or 550 kN. The maximum thrust may be in an inclusiverange bounded by any two of the values in the previous sentence (i.e.the values may form upper or lower bounds). The thrust referred to abovemay be the maximum net thrust at standard atmospheric conditions at sealevel plus 15 deg C. (ambient pressure 101.3 kPa, temperature 30 degC.), with the engine static.

In use, the temperature of the flow at the entry to the high-pressureturbine may be particularly high. This temperature, which may bereferred to as TET, may be measured at the exit to the combustor, forexample immediately upstream of the first turbine vane, which itself maybe referred to as a nozzle guide vane. At cruise, the TET may be atleast (or on the order of) any of the following: 1400K, 1450K, 1500K,1550K, 1600K or 1650K. The TET at cruise may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). The maximum TET in use of theengine may be, for example, at least (or on the order of) any of thefollowing: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. Themaximum TET may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds). The maximum TET may occur, for example, at a high thrustcondition, for example at a maximum take-off (MTO) condition.

A fan blade and/or aerofoil portion of a fan blade described herein maybe manufactured from any suitable material or combination of materials.For example, at least a part of the fan blade and/or aerofoil may bemanufactured at least in part from a composite, for example a metalmatrix composite and/or an organic matrix composite, such as carbonfibre. By way of further example at least a part of the fan blade and/oraerofoil may be manufactured at least in part from a metal, such as atitanium based metal or an aluminium based material (such as analuminium-lithium alloy) or a steel based material. The fan blade maycomprise at least two regions manufactured using different materials.For example, the fan blade may have a protective leading edge, which maybe manufactured using a material that is better able to resist impact(for example from birds, ice or other material) than the rest of theblade. Such a leading edge may, for example, be manufactured usingtitanium or a titanium-based alloy. Thus, purely by way of example, thefan blade may have a carbon-fibre or aluminium based body (such as analuminium lithium alloy) with a titanium leading edge.

A fan as described herein may comprise a central portion, from which thefan blades may extend, for example in a radial direction. The fan bladesmay be attached to the central portion in any desired manner. Forexample, each fan blade may comprise a fixture which may engage acorresponding slot in the hub (or disc). Purely by way of example, sucha fixture may be in the form of a dovetail that may slot into and/orengage a corresponding slot in the hub/disc in order to fix the fanblade to the hub/disc. By way of further example, the fan blades maybeformed integrally with a central portion. Such an arrangement may bereferred to as a blisk or a bling. Any suitable method may be used tomanufacture such a blisk or bling. For example, at least a part of thefan blades may be machined from a block and/or at least part of the fanblades may be attached to the hub/disc by welding, such as linearfriction welding.

The gas turbine engines described and/or claimed herein may or may notbe provided with a variable area nozzle (VAN). Such a variable areanozzle may allow the exit area of the bypass duct to be varied in use.The general principles of the present disclosure may apply to engineswith or without a VAN.

The fan of a gas turbine as described and/or claimed herein may have anydesired number of fan blades, for example 16, 18, 20, or 22 fan blades.

As used herein, cruise conditions may mean cruise conditions of anaircraft to which the gas turbine engine is attached. Such cruiseconditions may be conventionally defined as the conditions atmid-cruise, for example the conditions experienced by the aircraftand/or engine at the midpoint (in terms of time and/or distance) betweentop of climb and start of decent.

Purely by way of example, the forward speed at the cruise condition maybe any point in the range of from Mach 0.7 to 0.9, for example 0.75 to0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example0.78 to 0.82, for example 0.79 to for example on the order of Mach 0.8,on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Anysingle speed within these ranges may be the cruise condition. For someaircraft, the cruise conditions may be outside these ranges, for examplebelow Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond tostandard atmospheric conditions at an altitude that is in the range offrom 10000 m to 15000 m, for example in the range of from 10000 m to12000 m, for example in the range of from 10400 m to 11600 m (around38000 ft), for example in the range of from 10500 m to 11500 m, forexample in the range of from 10600 m to 11400 m, for example in therange of from 10700 m (around 35000 ft) to 11300 m, for example in therange of from 10800 m to 11200 m, for example in the range of from 10900m to 11100 m, for example on the order of 11000 m. The cruise conditionsmay correspond to standard atmospheric conditions at any given altitudein these ranges.

Purely by way of example, the cruise conditions may correspond to: aforward Mach number of 0.8; a pressure of 23000 Pa; and a temperature of−55 deg C.

As used anywhere herein, “cruise” or “cruise conditions” may mean theaerodynamic design point. Such an aerodynamic design point (or ADP) maycorrespond to the conditions (comprising, for example, one or more ofthe Mach Number, environmental conditions and thrust requirement) forwhich the fan is designed to operate. This may mean, for example, theconditions at which the fan (or gas turbine engine) is designed to haveoptimum efficiency.

In use, a gas turbine engine described and/or claimed herein may operateat the cruise conditions defined elsewhere herein. Such cruiseconditions may be determined by the cruise conditions (for example themid-cruise conditions) of an aircraft to which at least one (for example2 or 4) gas turbine engine may be mounted in order to provide propulsivethrust.

The skilled person will appreciate that except where mutually exclusive,a feature or parameter described in relation to any one of the aboveaspects may be applied to any other aspect. Furthermore, except wheremutually exclusive, any feature or parameter described herein may beapplied to any aspect and/or combined with any other feature orparameter described herein.

DESCRIPTION OF THE DRAWINGS

Embodiments will now be described by way of example only, with referenceto the Figures, in which:

FIG. 1 is a sectional side view of a gas turbine engine;

FIG. 2 is a close-up sectional side view of an upstream portion of a gasturbine engine;

FIG. 3 is a partially cut-away view of a gearbox for a gas turbineengine;

FIG. 4 is a sectional view of a prior art compressor assembly;

FIG. 5 is a sectional view of a first embodiment; and

FIG. 6 is a sectional view of a second embodiment.

DETAILED DESCRIPTION

FIG. 5 shows a first embodiment of a high-pressure compressor assemblycomprising a first compressor stage having a first disc portion 100 awith a peripheral first rim portion 101 a and, axially adjacent anddownstream of the first compressor stage, a second compressor stage, thesecond compressor stage having a second disc portion with a peripheralsecond rim portion (not shown).

The first rim portion (and the second rim portion) carries a series ofcircumferentially arranged compressor blades 102 a.

The second rim portion has an axial extension 104 b extending axiallyupstream towards the first rim portion 101 a. The first compressor stagecomprises a series of circumferentially arrange stators 103 which areradially aligned and radially outwards of the axial extension 104 of thesecond rim portion.

The axial extension 104 b defines a compressor drum cavity between thefirst and second disc portions. The second compressor stage furthercomprises a drive arm 106 extending from a radially inner surface of theaxial extension 104 b of the second rim portion within the drum cavityto the first disc portion 100 a. The drive arm 106 comprises a radialportion 106 a extending from an oblique portion 106 b.

The radial portion 106 b lies in abutment with the first disc portion100 a and a series of circumferentially arranged bolts 107 extendthrough the first disc portion 100 a and the radial portion 106 a of thedrive arm 106.

The oblique portion 106 b of the drive arm 106 divides the drum cavityinto radially outer rim cavity portion 108 and a radially inner maincavity portion 109.

In order to provide a pressure equalisation path from the rim cavityportion 108 to the main cavity portion 109, the oblique portion 106 b ofthe drive arm 106 comprises a plurality of circumferentially spacedapertures 111 extending from a radially outer surface in the rim cavityportion 108 to a radially inner surface in the main cavity portion 109.

The rim cavity 108 houses a rim seal 110 located between the axialextension 104 b of the second rim portion and the first rim portion 101a. The rim seal 110 comprises a folded metal sheet seal having a firstleaf 110 a in abutment with the first rim portion 101 a and a secondleaf 110 b in abutment a chamfered edge of with the axial extension 104b of the second rim portion.

The rim seal 110 and the apertures 111 which provide a pressureequalisation path between the rim cavity portion 108 and the main cavityportion 109, allow the pressure within the rim cavity portion 108 tosubstantially match the pressure in the main cavity portion 109 ratherthan the pressure of the gas path (which is radially outwards of the rimportions). Accordingly, the load applied to the bolts 107 issignificantly reduced because the lower pressure in the rim cavity 108no longer imparts a high axial separation pressure load on the bolts107. The bolts 107 only have to provide the structural connectionbetween the two compressor stages and no longer have to provide the sealagainst the pressure in the gas path.

FIG. 6 shows a second embodiment which is the same as the firstembodiment except that the rim seal comprises an interference seal 110and the pressure equalisation path is provided via the aperture 111′.

Other embodiments (not shown) may combine the rim seal of the firstembodiment with the pressure equalisation path of the second and viceversa.

The compressor assemblies described above are for use in a gas turbineengine such as that shown in FIG. 1 and discussed above.

Such a gas turbine engine 10 may comprise an engine core 11 comprisingat least one turbine 17, 19, a combustor 16, at least one compressor 14,15 which each comprise a compressor assembly as described above, and acore shaft 26. Such a gas turbine engine may comprise a fan 23 (havingfan blades) located upstream of the engine core 11.

Arrangements of the present disclosure may be particularly, although notexclusively, beneficial for fans 23 that are driven via a gearbox 30.Accordingly, the gas turbine engine may comprise a gearbox 30 thatreceives an input from the core shaft 26 and outputs drive to the fan 23so as to drive the fan 23 at a lower rotational speed than the coreshaft 26. The input to the gearbox 30 may be directly from the coreshaft 26, or indirectly from the core shaft 26, for example via a spurshaft and/or gear.

An exemplary arrangement for a geared fan gas turbine engine 10 is shownin FIG. 2 . The low-pressure turbine 19 (see FIG. 1 ) drives the shaft26, which is coupled to a sun wheel, or sun gear, 28 of the epicyclicgear arrangement 30. Radially outwardly of the sun gear 28 andintermeshing therewith is a plurality of planet gears 32 that arecoupled together by a planet carrier 34. The planet carrier 34constrains the planet gears 32 to precess around the sun gear 28 insynchronicity whilst enabling each planet gear 32 to rotate about itsown axis. The planet carrier 34 is coupled via linkages 36 to the fan 23in order to drive its rotation about the engine axis 9. Radiallyoutwardly of the planet gears 32 and intermeshing therewith is anannulus or ring gear 38 that is coupled, via linkages 40, to astationary supporting structure 24.

Note that the terms “low pressure turbine” and “low pressure compressor”as used herein may be taken to mean the lowest pressure turbine stagesand lowest pressure compressor stages (i.e. not including the fan 23)respectively and/or the turbine and compressor stages that are connectedtogether by the interconnecting shaft 26 with the lowest rotationalspeed in the engine (i.e. not including the gearbox output shaft thatdrives the fan 23). In some literature, the “low pressure turbine” and“low pressure compressor” referred to herein may alternatively be knownas the “intermediate pressure turbine” and “intermediate pressurecompressor”. Where such alternative nomenclature is used, the fan 23 maybe referred to as a first, or lowest pressure, compression stage.

The epicyclic gearbox 30 is shown by way of example in greater detail inFIG. 3 . Each of the sun gear 28, planet gears 32 and ring gear 38comprise teeth about their periphery to intermesh with the other gears.However, for clarity only exemplary portions of the teeth areillustrated in FIG. 3 . There are four planet gears 32 illustrated,although it will be apparent to the skilled reader that more or fewerplanet gears 32 may be provided within the scope of the disclosure.Practical applications of a planetary epicyclic gearbox 30 generallycomprise at least three planet gears 32.

The epicyclic gearbox 30 illustrated by way of example in FIGS. 2 and 3is of the planetary type, in that the planet carrier 34 is coupled to anoutput shaft via linkages 36, with the ring gear 38 fixed. However, anyother suitable type of epicyclic gearbox may be used. By way of furtherexample, the epicyclic gearbox 30 may be a star arrangement, in whichthe planet carrier 34 is held fixed, with the ring (or annulus) gear 38allowed to rotate. In such an arrangement the fan 23 is driven by thering gear 38. By way of further alternative example, the gearbox 30 maybe a differential gearbox in which the ring gear 38 and the planetcarrier 34 are both allowed to rotate.

It will be appreciated that the arrangement shown in FIGS. 2 and 3 is byway of example only, and various alternatives are within the scope ofthe present disclosure. Purely by way of example, any suitablearrangement may be used for locating the gearbox 30 in the engine 10and/or for connecting the gearbox 30 to the engine 10. By way of furtherexample, the connections (such as the linkages 36, 40 in the FIG. 2example) between the gearbox 30 and other parts of the engine 10 (suchas the input shaft 26, the output shaft and the fixed structure 24) mayhave any desired degree of stiffness or flexibility. By way of furtherexample, any suitable arrangement of the bearings between rotating andstationary parts of the engine (for example between the input and outputshafts from the gearbox and the fixed structures, such as the gearboxcasing) may be used, and the disclosure is not limited to the exemplaryarrangement of FIG. 2 . For example, where the gearbox 30 has a stararrangement (described above), the skilled person would readilyunderstand that the arrangement of output and support linkages andbearing locations would typically be different to that shown by way ofexample in FIG. 2 .

Accordingly, the present disclosure extends to a gas turbine enginehaving any arrangement of gearbox styles (for example star orplanetary), support structures, input and output shaft arrangement, andbearing locations.

Optionally, the gearbox may drive additional and/or alternativecomponents (e.g. the intermediate pressure compressor and/or a boostercompressor).

Other gas turbine engines to which the present disclosure may be appliedmay have alternative configurations. For example, such engines may havean alternative number of compressors and/or turbines and/or analternative number of interconnecting shafts. By way of further example,the gas turbine engine shown in FIG. 1 has a split flow nozzle 20, 22meaning that the flow through the bypass duct 22 has its own nozzle thatis separate to and radially outside the core engine nozzle 20. However,this is not limiting, and any aspect of the present disclosure may alsoapply to engines in which the flow through the bypass duct 22 and theflow through the core 11 are mixed, or combined, before (or upstream of)a single nozzle, which may be referred to as a mixed flow nozzle. One orboth nozzles (whether mixed or split flow) may have a fixed or variablearea. Whilst the described example relates to a turbofan engine, thedisclosure may apply, for example, to any type of gas turbine engine,such as an open rotor (in which the fan stage is not surrounded by anacelle) or turboprop engine, for example. In some arrangements, the gasturbine engine 10 may not comprise a gearbox 30.

The geometry of the gas turbine engine 10, and components thereof, isdefined by a conventional axis system, comprising an axial direction(which is aligned with the rotational axis 9), a radial direction (inthe bottom-to-top direction in FIG. 1 ), and a circumferential direction(perpendicular to the page in the FIG. 1 view). The axial, radial andcircumferential directions are mutually perpendicular.

It will be understood that the disclosure is not limited to theembodiments above-described and various modifications and improvementscan be made without departing from the concepts described herein. Exceptwhere mutually exclusive, any of the features may be employed separatelyor in combination with any other features and the disclosure extends toand includes all combinations and sub-combinations of one or morefeatures described herein.

We claim:
 1. A rotor assembly for a gas turbine engine, the rotorassembly comprising: a first rotor stage having a first disc portionwith a peripheral first rim portion; a second rotor stage, the secondrotor stage having a second disc portion with a peripheral second rimportion, wherein the second rotor stage is axially adjacent anddownstream of the first rotor stage and the second rim portion has anaxial extension extending towards the first rim portion such that aterminal end of the axial extension is adjacent the peripheral first rimportion, the axial extension of the second rim portion defining a rotordrum cavity between the first and second disc portions, the rotor drumcavity being located radially inward of the axial extension, wherein thesecond rotor stage further comprises a drive arm depending radiallyinward from the axial extension and extending radially inwardly withinthe drum cavity to the first disc portion, the drive arm being connectedto the first disc portion by at least one connector, wherein the drivearm divides the drum cavity into radially outer rim cavity portion and aradially inner main cavity portion, the radially outer rim cavityportion being located between a radially inwardly facing surface of theaxial extension and a radially outwardly facing surface of the drivearm, the at least one connector being located radially inwardly of theradially outer rim cavity portion and the drive arm and extending intothe radially inner main cavity portion, wherein the rotor assemblyfurther comprises a rim seal located between the axial extension of thesecond rim portion and the first rim portion, wherein the rotor assemblycomprises a pressure equalisation path extending from the rim cavityportion to the main cavity portion, wherein the drive arm comprises anangled portion and a radial portion, the angled portion extendingradially inwardly and axially forward from the radially inwardly facingsurface of the axial extension to the radial portion, and wherein theradial portion is located radially inward of and extending away from theangled portion, and is located axially forward of an axially aft-mostfacing surface of the peripheral rim portion of the first rotor stage,the axial aft-most facing surface of the peripheral rim portion beinglocated radially outwardly of the radial portion of the drive arm andbeing located adjacent to the terminal end of the axial extension ofsecond rotor stage.
 2. The assembly according to claim 1, wherein thepressure equalisation path is provided through the drive arm.
 3. Theassembly according to claim 1, wherein the angled portion of the drivearm comprises at least one aperture extending from the radiallyoutwardly facing surface in the rim cavity portion to a radially innersurface in the main cavity portion.
 4. The assembly according to claim1, wherein the rim seal is a folded sheet seal having a first leaf and asecond leaf, wherein the first leaf is in abutment with a first axiallyaft facing surface of the first rim portion located radially outwardlyof the drive arm, and wherein the second leaf is in abutment with theradially inner surface of the axial extension of the second rim portion.5. The assembly according to claim 1, wherein the rim seal is aninterference seal.
 6. The assembly according to claim 1, wherein therotor assembly is a compressor assembly and the first rotor stage is afirst compressor stage and the second rotor stage is a second, adjacentcompressor stage.
 7. A gas turbine engine comprising a rotor assemblyaccording to claim 1, wherein the rotor assembly is configured tooperate as a compressor assembly.
 8. The assembly of claim 1, whereinthe radial portion extends entirely radially along an axially aft facingsurface of the first disc portion.
 9. A rotor assembly for a gas turbineengine, the rotor assembly comprising: a first rotor stage having afirst disc portion with a peripheral first rim portion; a second rotorstage, the second rotor stage having a second disc portion with aperipheral second rim portion, the second rotor stage including aplurality of blades arranged circumferentially about the second discportion, wherein the second rotor stage is axially adjacent anddownstream of the first rotor stage and the second rim portion has anaxial extension extending towards the first rim portion such that aterminal end of the axial extension is adjacent the peripheral first rimportion, the axial extension defining a rotor drum cavity between thefirst and second disc portions, the rotor drum cavity being locatedradially inward of the axial extension, wherein the second rotor stagefurther comprises a drive arm coupled directly to the axial extensionand extending radially inward away from the axial extension andextending radially inwardly within the drum cavity to the first discportion, wherein an entirety of the drive arm is arranged axiallyforward of the plurality of blades of the second rotor stage, whereinthe drive arm is connected to the first disc portion by at least oneconnector, wherein the drive arm divides the drum cavity into radiallyouter rim cavity portion and a radially inner main cavity portion, theradially outer rim cavity portion being located between a radiallyinwardly facing surface of the axial extension and a radially outwardlyfacing surface of the drive arm, the at least one connector beinglocated radially inwardly of the radially outer rim cavity portion andthe drive arm and extending into the radially inner main cavity portion,wherein the rotor assembly further comprises a rim seal located betweenthe axial extension of the second rim portion and the first rim portion,wherein the rotor assembly further comprises a pressure equalisationpath extending from the rim cavity portion to the main cavity portion,wherein the at least one connector is a bolt extending through the firstrotor stage and the drive arm so as to couple the second rotor stage tothe first rotor stage, and wherein the terminal end of the axialextension is located axially forward of at least a portion of the atleast one bolt.
 10. The assembly according to claim 9, wherein thepressure equalisation path is provided through the drive arm.
 11. Theassembly according to claim 9, wherein an angled portion of the drivearm comprises at least one aperture extending from a radially outersurface in the rim cavity portion to a radially inner surface in themain cavity portion.
 12. The assembly according to claim 9, wherein therim seal is a folded sheet seal having a first leaf and a second leaf,wherein the first leaf is in abutment with a first axially aft facingsurface of the first rim portion located radially outwardly of the drivearm, and wherein the second leaf is in abutment with the radially innersurface of the axial extension of the second rim portion.
 13. Theassembly according to claim 9, wherein the drive arm further comprises aradial portion extending radially inwardly of an angled portion, andwherein the radial portion extends entirely radially along an axiallyaft facing surface of the first disc portion.
 14. The assembly accordingto claim 9, wherein the rotor assembly is a compressor assembly and thefirst rotor stage is a first compressor stage and the second rotor stageis a second, adjacent compressor stage.
 15. A gas turbine engine,comprising: a first rotor stage having a first disc portion with aperipheral first rim portion; a second rotor stage, the second rotorstage having a second disc portion with a peripheral second rim portion,the second rotor stage including a plurality of blades arrangedcircumferentially about the second disc portion; and a vane stagearranged axially between the first and second rotor stages and includingat least one vane, wherein the second rotor stage is axially adjacentand downstream of the first rotor stage and the second rim portion hasan axial extension extending towards the first rim portion such that aterminal end of the axial extension is adjacent the peripheral first rimportion, the terminal end of the axial extension being located axiallyforward of the at least one vane of the vane stage, the axial extensiondefining a rotor drum cavity between the first and second disc portions,the rotor drum cavity being located radially inward of the axialextension, wherein the second rotor stage further comprises a drive armcoupled directly to the axial extension and extending radially inwardaway from the axial extension and extending radially inwardly within thedrum cavity to the first disc portion, wherein an entirety of the drivearm is arranged axially forward of the plurality of blades of the secondrotor stage, wherein the drive arm is connected to the first discportion by at least one connector, wherein the drive arm divides thedrum cavity into radially outer rim cavity portion and a radially innermain cavity portion, the radially outer rim cavity portion being locatedbetween a radially inwardly facing surface of the axial extension and aradially outwardly facing surface of the drive arm, the at least oneconnector being located radially inwardly of the radially outer rimcavity portion and the drive arm and extending into the radially innermain cavity portion, wherein the rotor assembly further comprises a rimseal located between the axial extension of the second rim portion andthe first rim portion, and wherein the rotor assembly comprises apressure equalisation path extending from the rim cavity portion to themain cavity portion.
 16. The gas turbine engine according to claim 15,wherein the rim seal is a folded sheet seal having a first leaf and asecond leaf, wherein the first leaf is in abutment with a first axiallyaft facing surface of the first rim portion located radially outwardlyof the drive arm, and wherein the second leaf is in abutment with theradially inner surface of the axial extension of the second rim portion.